1. Field
This invention relates to a propulsion system for a missile. More particularly, the invention relates to a system for separating and controlling oxidizer and fuel in an insensitive missile propulsion system, thereby enabling efficient operation in each one of the multiple modes of operation utilized by the missile during a flight.
2. Description of the Related Art
Conventional rocket propulsion systems to propel missiles have performance limitations imposed by a number of factors including a requirement to transport a required amount of oxidizer. For a given take-off mass this results in shorter range powered flights or reduced payloads relative to systems like ramjets and scramjets which obtain the oxidizer from atmospheric air. Ramjet and scramjet engines have their own limitations; including inadequate thrust at low speeds thereby requiring a rocket or turbine booster of significant mass to accelerate the missile to ramjet takeover speed. Further, since the oxidizer for the ramjet comes from the atmosphere, the ramjet fuel flow must be controlled during flight to maintain the proper fuel to oxidizer ratio which may vary greatly over the flight duration.
Variable cycle engines that transition from a rocket propulsion system at launch and relatively low speeds to a ramjet propulsion system at higher speeds are known from U.S. Pat. Nos. 4,651,523 and 5,224,344, both of which are incorporated by reference in their entireties herein. U.S. Pat. No. 4,651,523 discloses a dual cycle engine having a solid propellant with an aft end shaped to form a nozzle. As the propellant burns, the missile is propelled by rocket thrust. Burning of the propellant allows forwardly positioned air covers to slide rearward enabling compressed air to be delivered to a combustion chamber initiating ramjet operation. U.S. Pat. No. 5,224,344 discloses a rocket chamber replacing the ramjet pilot of a Dual Combustion Ramjet (DCR) engine. A mixture of liquid fuel and oxidizer combusted within the rocket engine initially propels the missile by rocket propulsion. As the speed increases, compressed air is delivered to the ramjet combustor through inlet passages and additional fuel is provided to the rocket engine delivering a fuel-rich exhaust to the combustor. Combustion of the fuel-rich exhaust heats and expands the compressed air enabling ramjet operation.
Most missiles today employ solid rocket propellants that contain an intimate mixture of fuel and oxidizer chemicals which when ignited produce a highly energetic stream of gas used effectively for propulsion. Liquid bi-propellant rocket systems separate the fuel and oxidizer until injected into the rocket motor but have a risk of leakage and fire if the tanks are breached by accident or enemy action. Accidental ignition of these propellant(s) can cause severe hazards for the user of the weapon. These hazards include explosion, fire or even uncommanded flight of the missile. The US Defense Department is demanding missile propulsion systems that are insensitive to accidental ignition that may occur during handling or at any time prior to the planned launch of the weapon. Systems that have reduced hazards are referred to as Insensitive Munitions or IM.
Hybrid engines having a solid component and a liquid component are one type of IM. When the fuel is liquid and the oxidizer solid, the engine is referred to as a reverse hybrid engine. One reverse hybrid engine is disclosed in U.S. Pat. No. 3,555,826, which is incorporated by reference in its entirety herein. The patent discloses an engine having the liquid fuel separated from a solid oxidizer by an electrically actuated mechanical valve.
There remains a need for an effective variable cycle engine capable of utilizing insensitive munitions for propulsion.